Turbomachine aircraft propeller

ABSTRACT

The invention relates to an aircraft propeller ( 1 ) comprising a turbomachine ( 8 ) housed in a nacelle ( 10 ) and a cooler ( 14 ) capable of being traversed by a hot fluid, which is to be cooled by thermal exchange with cold air external to the cooler. The propeller ( 1 ) comprises an air vein ( 13 ) ( 13   b ) capable of directing pressurized air towards an air duct ( 20 ) realized between an outer wall ( 6 ) and an inner wall ( 60 ) of the nacelle ( 10 ). The cooler ( 14 ) comprises a first cooling means, called first surface cooling means ( 145 ), on a first surface ( 141 ), arranged at the outer wall ( 6 ) of the propeller nacelle ( 10 ) and a second surface cooling means ( 146 ), on a second surface ( 142 ) arranged at a wall ( 23 ) of the air duct ( 20 ). The invention also concerns an aircraft equipped with such a propeller.

The present invention relates to a turbomachine-type aircraft propeller. More specifically, the invention relates to a cooling device for propellers.

In aviation, a large number of aircraft propellers comprise a turbomachine, housed in a nacelle.

One can cite the case, for example, of turbofan-type propellers for which the turbomachine drives at least one rotor located inside the nacelle. One can also cite the case of “propfan” or “open rotor” type propellers, for which the turbomachine drives two counter-rotating rotors, the two rotors being located outside the nacelle, downstream or upstream of the turbomachine.

Whatever the type of propellers, a gear box (gears between the turbomachine shaft and the rotors) transmits the mechanical energy generated by the turbomachine to the rotors.

Although it has very high efficiency, this gearbox dissipates part of the energy created by the propeller into heat by friction. This heat is transmitted in particular to the gearbox lubrication fluid.

Moreover, the turbomachine itself generates significant heat dissipation mainly by mechanical friction, also through its lubricating fluid.

It is clear that this heat must be dissipated to the outside environment to cool the propeller.

Equipment mounted on the turbomachine, such as an electric generator, may also require cooling.

Various solutions have been developed to perform this cooling.

A first known solution, mainly for turbofan-type propellers, is to install a heat exchanger, known as a “volumetric heat exchanger”, between an outer wall and an inner wall of the nacelle. An air inlet collects cold air from the cold air flow going through the turbomachine, to bring it inside said volumetric heat exchanger. After passing through the heat exchanger matrix, the air is ejected out of the nacelle through an air outlet. Such heat exchangers have not proved to be an optimal solution in terms of propulsive efficiency and of aerodynamic impact on the turbomachine. In effect, the air collection represents a direct loss of propulsive efficiency inasmuch as it contributes little or nothing to the engine's thrust. Moreover, the presence of an air inlet, one or more internal ducts and an air outlet generates load losses and disturbs the propeller's internal flow more or less significantly.

Another known solution is to use an exchanger, known as a “surface exchanger”, e.g. a plate heat exchanger. In particular, a surface exchanger is known that locally takes the shape of an inner wall of the nacelle or of the engine cover to which it is coupled. A first surface of the surface exchanger is coupled to the inner wall of the nacelle or to the engine cover, while a second surface is located in the flow of cold air flowing through the internal volume of the nacelle. The heat transported within the heat exchanger is transferred by thermal conduction to the inner surface of the plate forming the lower surface of the plate heat exchanger. This hot plate is traversed by the flow of cold air flowing in the nacelle. The heat stored in the hot plate on the inner surface is thus dissipated by forced convection towards the propeller's airflow.

This solution still has an aerodynamic impact, but has the advantage, compared to the previous solution, of not collecting air from the flow through the turbomachine.

However, this solution cannot be transposed to propfan-type propellers. Indeed, when the aircraft speed is low or zero, there is little or no air flow traversing the surface exchanger, because the rotors are outside the nacelle.

In the case of a propfan propeller, the surface heat exchanger can only be positioned on the outer wall of the nacelle or on the propeller pylon.

The objective of this invention is therefore to provide an propeller comprising a turbomachine cooling device that overcomes the aforementioned drawbacks by ensuring an adequate level of cooling on the ground and in flight while limiting the aerodynamic impact during flight phases.

To this end, the invention envisages an aircraft propeller comprising a turbomachine housed in a nacelle and a cooler capable of being traversed by a hot fluid, which is to be cooled by thermal exchange with cold air external to the cooler. The propeller comprises at least one air vein capable of directing pressurized air towards an air duct realized between an outer wall and an inner wall of the nacelle, and the cooler comprises:

-   -   a first cooling means, called first surface cooling means, on a         first surface, called outer surface, arranged at the outer wall         of the propeller nacelle,     -   a second cooling means, called second surface cooling means, on         a second surface, called inner surface, arranged at an inner         wall of the air duct.

In other words, the cooler is intended to operate during all the phases of piloting an aircraft firstly, when the aircraft is at zero or low speed and secondly, when the aircraft is in flight.

Preferably, the first surface cooling means is sized so as to be sufficient to ensure the desired cooling when the aircraft is in flight, within preselected environmental and speed conditions, and the second surface cooling means is sized so as to be sufficient to ensure the desired cooling when the aircraft is at low or zero speed, within preselected environmental conditions.

Throughout the description, “upstream” shall designate, at a given point, the part that is placed in front of this point by reference to the direction of the airflow in the propeller, and “downstream” shall designate the part that is located behind this point.

In a first embodiment of the invention, an air inlet is located at a 1^(st) stage of an air compressor of the turbomachine and the associated air vein emerges upstream of cooler.

According to another embodiment of the invention, a pressurized air inlet is located downstream from an air compressor of the turbomachine and the associated air vein emerges downstream from the cooler, to create suction at through it.

Advantageously, the air vein comprises, at the air duct 20, a small diameter pipe or ejector, which ejects pressurized air into said air duct.

According to another embodiment of the invention, the propeller comprises two air veins and a regulator valve associated with each vein, an air vein emerging upstream of the cooler and an air vein emerging downstream from the cooler, the regulator valves controlling the entry of pressurized air according to one vein or the other.

In one embodiment of the cooler, at least one of the first and second surface cooling means is a set of fins extending from the outer surface, and oriented mostly parallel to the direction of the airflow.

In another embodiment, the first and second surface cooling means have a similar architecture.

Alternatively, the air duct comprises an air inlet installed at the front of the propeller to create an additional air inlet.

The invention also envisages an aircraft comprising an propeller as set forth.

The description that will follow, given solely as an example of an embodiment of the invention, is made with reference to the figures included in an appendix, in which:

FIG. 1 shows an propeller of a type called “propfan”, to which the invention can be applied,

FIG. 2 illustrates such an propeller in a very schematic cross-section view,

FIG. 3 is a detail view of FIG. 2, centered on the front part of the propeller, which highlights the main elements of the cooling device according to a first embodiment of the invention,

FIG. 4 illustrates schematically the data processed by the electronic control of the cooling device according to the invention,

FIG. 5 a shows in a detail view the airflow in the cooler when the aircraft is at low speed, for the first embodiment of the invention,

FIG. 5 b shows in a detail view the airflow in the cooler when the aircraft is in the air, for the first embodiment of the invention,

FIG. 6 a shows in a detail view the airflow in the cooler when the aircraft is at low speed, according to a variant of the first embodiment of the invention,

FIG. 6 b shows in a detail view the airflow in the cooler when the aircraft is in flight, according to a variant of the first embodiment of the invention,

FIG. 7 a is a detail view of FIG. 2, centered on the front part of the propeller, which highlights the main elements of the cooling device according to a second embodiment of the invention,

FIG. 7 b shows in a detail view the airflow in the cooler when the aircraft is at low speed, for the second embodiment of the invention,

FIG. 7 c shows in a detail view the airflow in the cooler when the aircraft is in the air, for the second embodiment of the invention,

FIG. 8 a is a detail view of FIG. 2, centered on the front part of the propeller, which highlights the main elements of the cooling device according to a third embodiment of the invention,

FIG. 8 b shows in a detail view a first example of airflow in the cooler when the aircraft is at low speed, for the third embodiment of the invention,

FIG. 8 c shows in a detail view a second example of airflow in the cooler when the aircraft is at low speed, for the third embodiment of the invention,

FIG. 8 d shows in a detail view the airflow in the cooler when the aircraft is in the air, for the third embodiment of the invention,

FIG. 9 a shows in a detail view the airflow in the cooler when the aircraft is at low speed, according to a variant of the third embodiment of the invention,

FIG. 9 b shows in a detail view the airflow in the cooler when the aircraft is in flight, according to a variant of the third embodiment of the invention.

The invention relates to an aircraft propeller 1, for example of the type called “propfan” as shown in FIG. 1. Such propellers are envisaged for future aircraft. In the example of implementation illustrated here, two propfan propellers 1, each housed in a nacelle 10, are attached by propeller pylons, on both sides of an aircraft fuselage 2.

Each propfan propeller 1 comprises here two counter-rotating rotors 3 a, 3 b each comprising a set of blades 4 a, 4 b, which are equidistant and arranged at the rear of the propeller 1. The blades 4 a, 4 b of each rotor 3 a, 3 b protrude from an annular crown 5 a, 5 b, which is mobile with this rotor, an outer surface of which is located in the continuity of an outer wall 6 of the propeller nacelle.

As shown schematically in FIG. 2 the propfan propeller 1 comprises an air inlet 7 that supplies a turbomachine 8. This turbomachine 8 comprises an axial portion driven in rotation when the turbomachine is running. In turn, this shaft drives the shafts 9 a, 9 b of the blades 4 a, 4 b of the two counter-rotating rotors 3 a, 3 b via mechanical transmissions not shown in FIG. 2.

The hot gases generated by the turbomachine 8 when in operation are discharged through an annular hot vein 18 having an outlet located at the rear of the two rotors 3 a, 3 b. In a variant, these gases can also be discharged upstream of the two rotors.

The turbomachine 8 comprises, conventionally, a multistage compressor allowing incremental increases in the pressure of air entering the turbomachine.

The construction details of propfan propellers and their components—rotors, turbomachine, transmission—as well as their dimensions, materials etc. are beyond the scope of the present invention. The elements described here are therefore provided only for information purposes, to facilitate understanding of the invention in one of its non-limiting examples of implementation.

During the aircraft's flight, the outside air, whose temperature is between +55° C. near the ground and −74° C. at altitude, circulates along the outer wall 6 of the propeller's nacelle 10, substantially in the direction opposite to a longitudinal axis X of movement of the aircraft.

At the same time, the propeller generates a significant heat rejection, part of which is discharged through the annular hot duct 18, and another part, which is transmitted to the engine and gearbox oil circuits, must be discharged by an appropriate cooling device.

GENERAL DESCRIPTION

The cooling device comprises a cooler 14 between the outer wall 6 and an air duct 20 realized between the outer wall 6 and an inner wall 60 of the nacelle 10.

As illustrated in FIGS. 2 and 3 for example, the inner wall 60 is the wall located across from the turbomachine, in an internal volume of the nacelle 10 and the outer wall 6 is the wall swept by the flow of outside air.

In an embodiment of the air duct, as shown in FIG. 3, the air duct 20 firstly emerges at the front of the propeller nacelle by an air inlet 21, near the main air inlet 7 and secondly emerges outside the nacelle by an air outlet 22, upstream of the rotors.

The cooler 14 is designed to operate in two main modes of heat exchange:

-   -   one on the ground or during take-off, when the outside air flow         is low or zero,     -   the other in flight, when the flow of outside air is         significant.

The cooler 14 comprises a first surface cooling means 145 on a first surface, called outer surface 141, arranged at the outer wall 6 of the nacelle 10 of the propeller. It comprises a second surface cooling means 146 on a second side opposite the outer surface, called inner face 142, arranged at a wall 23 of the air duct 20.

The first surface cooling means 145 thus forms part of the outer wall 6 of the propeller nacelle 10. The shape of the first surface cooling means 145 is determined by the shape of the outer wall 6 of the nacelle 10 from the propeller at the place where the cooler 14 must be installed.

The second surface cooling means 146 forms part of the wall 23 of the air duct 20. The shape of the second surface cooling means 146 is determined by the shape of the wall 23 of the air duct at the place where the cooler 14 must be installed.

The inner surface 142 of the cooler 14 is, in this non-limiting example of the invention, substantially parallel to the outer surface 141.

The dimensions of the first 145, respectively second 146, surface cooling means are determined by the cooling requirement when the aircraft is in flight, respectively on the ground or at low speeds, by the flow of pressurized air cooling available.

The calculation itself is known to the man skilled in the art and is therefore not detailed further here.

In an example of realization of the first and/or second surface cooling means, the first surface cooling means comprises a set of fins (not shown in the figures) extending from the outer surface of the cooler and protruding on the outer surface of the cooler.

In another example of realization of the first and/or second surface cooling means, the second surface cooling means comprises a set of fins (not shown in the figures) extending from the inner surface of the cooler and protruding on the inner surface of the cooler.

For example, these fins can increase the exchange area, and are oriented substantially parallel to the flow lines of an air stream flowing over the outer (inner) surface of the cooler when the aircraft is in flight (at low speed), i.e. substantially along the longitudinal axis X.

The dimensions of these fins are determined by the cooling requirement when the aircraft is in flight or at low speed, and by the external air flow and the temperature of the air flowing along the surface of these fins. The details of such a calculation are known to the man skilled in the art.

These cooling means 145, 146 are known to the man skilled in the art and will not be developed further here.

In one embodiment of the cooler 14, said first and second surface cooling means are identical.

In a preferred embodiment of the cooler, the first cooling means 145 does not comprise fins on the outer face 141 of the cooler 14 and the second surface cooling means 145 has fins on the inner surface 142 of the cooler 14.

The cooling device is controlled by an electronic control unit 19 (shown in FIG. 4), of a type known per se.

In this non-limiting example, said electronic control unit 19 receives oil circuit temperature data that the cooling device must regulate, as well as outside air temperature data as inputs.

Said electronic control unit 19 transmits control data, e.g. temperature of the oil circuits, to the aircraft's cockpit, from which it also receives instructions.

This electronic control unit 19 may be installed at the propeller, close to the cooler 14. Alternatively, the electronic control unit 19 may be part of the various pieces of electronic equipment located in the cockpit, or simply be one of the functions provided by a multi-purpose computer usually found in aircraft.

In a variant of embodiment of the invention, the air duct 20 comprises, located at the air outlet 22, means of closing 30 said outlet. This means of closing is set by the electronic control unit 19.

In an example of realization, the means of closing 30 is a valve.

FIRST EMBODIMENT

In a first embodiment of the cooling device, said cooling device, as shown in FIG. 3, takes advantage of the presence of the compressor, and comprises an air inlet 11 of a type known per se, arranged in this non-limiting example, at a first stage of the compressor of the turbomachine 8. This arrangement is intended to provide air that is as yet little warmed by compression, instead of the air located at the following stages of the compressor.

The position of the collection point naturally depends on the specific characteristics of the turbomachine 8 under consideration and of its compressor, but this position is imposed by the requirement for air at a pressure sufficient to bring a predefined flow of air to a cooler at a sufficiently low temperature, while not disturbing the correct operation of the compressor and more generally of the turbomachine 8.

Preferably, this air inlet 11 comprises a regulator valve 12, here illustrated schematically, designed to control the flow of pressurized air collected at the air inlet 11 from a value close to zero to a maximum value determined by the cooling requirement of the gearbox and/or engine and/or electrical generator oil.

An air vein 13 located downstream from the regulator valve 12 directs the flow of pressurized air collected upstream of the cooler 14 towards the air duct 20.

The electronic control unit 19 sets the regulator valve 12 according to various input information. It receives temperature data from air in the air vein 13 and regulator valve 12 status information.

In operation, when the aircraft is on the ground (FIG. 5 a) or in taxiing, takeoff or approach phases, with the propellers operating, the thermal discharge from the electrical generator is very large and the aircraft speed is low or zero.

During these phases, called low speeds phases, the flow of outside air is low and insufficient for cooling by the two surface cooling means 145, 146. The electronic control unit 19 sets the regulator valve 12 substantially into the maximum open position, allowing the second cooling means 146 to be traversed by outside air and pressurized air collected at the compressor. The cooling is performed mainly by the second cooling means 146.

This ensures a heat exchange between the hot cooling means 14 and the cold pressurized air, causing the desired cooling of the cooler 14 and of the fluids circulating within or connected to it by thermal conduction.

As the climb progresses and evolves towards level flight, the speed of the aircraft increases and the outside air temperature decreases. Accordingly, the collection of air at the compressor is reduced by gradual closing of the regulator valve 12 controlled by the electronic control unit 19 and the cooling is performed increasingly firstly by the first surface cooling means 145 traversed by the outside air and secondly by the second surface cooling means 146 traversed by the outside air flowing naturally into the air duct 20.

The closing (and by extension, the opening) of the valve 12 is described to be gradual but it is also possible that the closing (and by extension, the opening) is controlled in an on-or-off manner.

Subsequently, when the aircraft is in steady flight (FIG. 5 b), the cooling is performed normally by the first and second cooling means 145, 146 of the cooler 14, mainly by the first cooling means 145, and the regulator valve 12 then remains closed, thereby eliminating the air collection from the compressor, and therefore reducing the increased fuel consumption that otherwise arises from this power draw.

In variant of realization, when the air duct 20 comprises a means of closing 30, the electronic control unit 19 preferably sets the means of closing into closed position during flight phases. In closed position, the means of closing 30 limits the impact of aerodynamic drag.

In variant of realization, as shown in FIGS. 6 a and 6 b, the air duct 20 does not emerge, for example using means of closing air inlet, toward the front of the nacelle 10, so as to reduce the aerodynamic drag caused by the air inlet.

During the low speed phases (FIG. 6 a), the electronic control unit 19 sets the regulator valve 12 substantially into the maximum open position and pressurized cold air flows through the air duct 20. The first and second cooling means are in operation. When the air duct 20 further comprises a means of closing of the air outlet 22, said means of closing is in the open position.

During the flight phase (FIG. 6 b), the electronic control unit 19 sets the regulator valve 12 into the closed position and the cooling is performed only by the first surface cooling means 145. When the air duct 20 further comprises a means of closing of the air outlet 22, said means of closing is preferably in the closed position.

SECOND EMBODIMENT

In a second embodiment of the cooling device, said cooling device, as shown in FIGS. 7 a to 7 c comprises an air inlet 11 b of a type known per se, arranged in this non-limiting example, upstream of the compressor of the turbomachine 8.

Preferably, this air inlet 11 b comprises a regulator valve 12 b, here illustrated schematically, designed to control the flow of pressurized air collected at the air inlet 11 b from a value close to zero to a maximum value determined by the cooling requirement of the gearbox and/or engine and/or electrical generator oil.

An air vein 13 b located downstream from the regulator valve 12 b directs the flow of pressurized air collected towards the air duct 20 realized in the nacelle, downstream from the cooler 14, and creates suction of outside air from the air inlet 21 into the air duct, traversing the second cooling means 146.

Advantageously, the air vein 13 b ends, at the air duct 20, in a small diameter pipe or ejector, which ejects pressurized air into the air duct 20. The ejection of the pressurized air produces an acceleration of the external airflow coming from the air duct 20 by a suction phenomenon and consequently an increase in the air flow traversing the second cooling means 146.

In this second embodiment, the electronic control unit 19 sets the regulator valve 12 b according to various input information. It receives temperature data from air in the air vein 13 b and regulator valve 12 b status information.

In operation, when the aircraft is in low speed phases (FIG. 7 b), the thermal discharge from the electrical generator is very large and the aircraft speed is low or zero.

During these low speeds phases, the flow of outside air is low and insufficient for cooling by the two surface cooling means 145, 146. The electronic controller 19 sets the regulator valve 12 b substantially in the maximum open position, to create suction of outside air at the exit of the air duct. The cooling is performed mainly by the second cooling means 146.

This ensures a heat exchange between the hot cooling means 14 and the cold pressurized air, causing the desired cooling of the cooler 14 and of the fluids circulating within or connected to it by thermal conduction.

As the climb progresses and evolves towards level flight, the speed of the aircraft increases and the outside air temperature decreases. Accordingly, the collection of air at the compressor is reduced by gradual closing of the regulator valve 12 b controlled by the electronic control unit 19 and the cooling is performed increasingly firstly by the first surface cooling means 145 traversed by the outside air and secondly by the second surface cooling means 146 traversed by the outside air flowing naturally into the air duct 20.

The closing (and by extension, the opening) of the valve 12 b is described to be gradual but it is also possible that the closing (and by extension, the opening) is controlled in an on-or-off manner.

Subsequently, when the aircraft is in steady flight (FIG. 7 c), the cooling is performed normally by the first and second cooling means 145, 146 of the cooler 14, mainly by the first cooling means 145, and the regulator valve 12 b then remains closed, thereby eliminating the air collection from the compressor, and therefore reducing the increased fuel consumption that otherwise arises from this power draw.

In variant of realization, when the air duct 20 comprises a means of closing 30, the electronic control unit 19 preferably sets the means of closing into closed position during flight phases. In closed position, the means of closing 30 limits the impact of aerodynamic drag.

THIRD EMBODIMENT

In a third embodiment of the cooling device, the cooling device, as shown in FIGS. 8 a to 8 d, comprises the air inlet 11, the regulator valve 12 and the air vein 13, such as described in the first embodiment.

The cooling device further comprises the air inlet 11 b, the regulator valve 12 b and the air vein 3 b as described in the second embodiment.

Advantageously, the air vein 13 b ends, at the air duct 20, in an ejector, which ejects pressurized air into the air duct 20.

In this third embodiment, the electronic control unit 19 sets the regulator valves 12 and 12 b according to various input information.

In operation, when the aircraft is in low speed phases (FIGS. 8 b and 8 c), the thermal discharge from the electrical generator is very large and the aircraft speed is low or zero.

During these low speeds phases, the flow of outside air is low and insufficient for cooling by the two surface cooling means 145, 146. The electronic control unit 19 therefore sets one of the two regulator valves 12, 12 b, substantially into the maximum open position, allowing either the second cooling means 146 to be traversed by outside air and pressurized air collected at the compressor when the regulator valve 12 is in the open position (FIG. 8 b), or to create suction of outside air at the exit of the air duct when the regulator valve 12 b is in the open position (FIG. 8 c). The cooling is performed mainly by the second cooling means 146.

This ensures a heat exchange between the hot cooling means 14 and the cold pressurized air, causing the desired cooling of the cooler 14 and of the fluids circulating within or connected to it by thermal conduction.

As the climb progresses and evolves towards level flight, the speed of the aircraft increases and the outside air temperature decreases. Accordingly, the collection of air at the compressor is reduced by gradually closing the regulator valve 12 or 12 b controlled by the electronic control unit 19 and the cooling is performed increasingly firstly by the surface cooling means 145 traversed by the outside air and secondly by the surface cooling means 146 traversed by the outside air flowing naturally into the air duct 20.

The closing (and by extension, the opening) of the valves 12,12 b is described as being gradual but it is also possible that the closing (and by extension, the opening) of the valves is controlled in an on-or-off manner.

Subsequently, when the aircraft is in steady flight (FIG. 8 d), the cooling is performed normally by the first and second cooling means 145, 146 of the cooler 14, mainly by the first cooling means 145, and the regulator valves 12 and 12 b then remains closed, thereby eliminating the air collection from the compressor, and therefore reducing the increased fuel consumption that otherwise arises from this power draw.

In variant of realization, when the air duct 20 comprises a means of closing 30, the electronic control unit 19 preferably sets the means of closing into closed position during flight phases. In closed position, the means of closing 30 limits the impact of aerodynamic drag.

In a variant of realization, as shown in FIGS. 9 a and 9 b, the air duct 20 does not emerge, for example using means of closing air inlet 21, toward the front of the nacelle 10, so as to reduce the aerodynamic drag caused by the air inlet.

In this embodiment variant, the regulator valve 12 b is always in the closed position, during both the low speed phases or in flight.

During the low speed phases (FIG. 9 a), the electronic control unit 19 sets the regulator valve 12 substantially into the maximum open position and pressurized cold air flows through the air duct 20. The first and second cooling means are in operation. When the air duct 20 further comprises a means of closing 30 of the air outlet 22, said means of closing is in the open position.

During the flight phase (FIG. 9 b), the electronic control unit 19 sets the regulator valve 12 into the closed position and the cooling is performed only by the first surface cooling means 145. When the air duct 20 further comprises a means of closing 30 of the air outlet 22, said means of closing is preferably in the closed position.

The scope of this invention is not limited to the details of the forms of embodiment considered above as an example, but on the contrary extends to modifications in the reach of the man skilled in the art.

The invention is described in the case of a propfan-type propeller, but the invention is also applicable to turbofan-type propellers.

It is apparent from the description that the cooling device allows the engine components to be cooled across all flight phases while allowing the use of a surface exchanger during low speed phases.

The fact of managing the opening and closing of the regulator valves 12 and/or 12 b during the low speed phase and in-flight allows the power draw on the compressor to be controlled, and to reduce it whenever possible, which translates into reduced consumption.

In addition, the present invention takes advantage of the presence of a surface exchanger, which is less than 6 cm thick, instead and in place of a volumetric exchanger, which is thicker than 15 cm. This results in a smaller footprint, lower air duct height, which limits the impact of aerodynamic drag on the one hand and facilitates its integration into propeller propellant on the other. 

1. Aircraft propeller (1) comprising a turbomachine (8) housed in a nacelle (10) and a cooler (14) capable of being traversed by a hot fluid, which is to be cooled by thermal exchange with cold air external to the cooler, the propeller (1) comprising an air vein (13) (13 b) capable of directing pressurized air towards an air duct (20) realized between an outer wall (6) and an inner wall (60) of the nacelle (10), the cooler (14) comprising: a first cooling means, called first surface cooling means (145), on a first surface, called outer surface (141), arranged at the outer wall (6) of the propeller (1) nacelle (10), a second cooling means, called second surface cooling means, on a second surface (146), called inner surface (142), arranged at a wall (23) of the air duct (20), wherein the air duct (20) comprises an air inlet (21) arranged at the front of the propeller to create an additional air inlet.
 2. Aircraft propeller according to claim 1, wherein: the first surface cooling means (145) is sized so as to be sufficient to ensure the desired cooling when the aircraft is in flight, within preselected environmental and speed conditions, the second surface cooling means (146) is sized so as to be sufficient to ensure the desired cooling when the aircraft is at low or zero speed, within preselected environmental conditions.
 3. Aircraft propeller according to any one of the preceding claims, wherein an air inlet (11) is located at a 1^(st) stage of an air compressor of the turbomachine (8) and in that the associated air vein (13) emerges upstream of cooler (14).
 4. Aircraft propeller according to any one of claims 1 to 2, wherein an air inlet (11 b) is located downstream from an air compressor of the turbomachine (8) and in that the associated air vein (13 b) emerges downstream from the cooler (14).
 5. Aircraft propeller according to claim 4, wherein the air vein (13 b) comprises, at the air duct (20), an ejector, which ejects pressurized air into said air duct (20).
 6. Aircraft propeller according to one of claims 1 to 2, comprising two air veins (13, 13 b) and a valve (12, 12 b) associated with each vein, an air vein (13) emerging upstream of the cooler (14) and an air vein (13 b) emerging downstream from the cooler, the valves (12,12 b) controlling the entry of pressurized air according to one vein or the other.
 7. Aircraft propeller according to any one of the preceding claims, wherein at least one of the first (145) and second (146) surface cooling means is a set of fins extending from the outer/inner surface, and oriented mostly parallel to the direction of the airflow.
 8. Aircraft propeller according to any one of the preceding claims, wherein the first and second surface cooling means (145, 146) are of similar architecture.
 9. Aircraft comprising a propeller according to any one of the preceding claims. 